Method for producing a high strength Al-Zn-Mg-Cu alloy

ABSTRACT

The present invention relates to a method for producing a high strength Al—Zn—Cu—Mg alloy with an improved fatigue crack growth resistance and a high damage tolerance, comprising the steps of casting an ingot with the following composition (in weight percent) Zn 5.5-9.5, Cu 1.5-3.5, Mg 1.5-3.5, Mn&lt;0.25, Zr&lt;0.25, Cr&lt;0.10, Fe&lt;0.25, Si&lt;0.25, Ti&lt;0.10, Hf and/or V&lt;0.25, other elements each less than 0.05 and less than 0.15 in total, balance aluminum, homogenizing and/or pre-heating the ingot after casting, hot working the ingot and optionally cold working into a worked product of more than 50 mm thickness, solution heat treating, quenching the heat treated product, and artificially ageing the worked and heat-treated product, wherein the ageing step comprises a first heat treatment at a temperature in a range of 105° C. to 135° C. for more than 2 hours and less than 8 hours and a second heat treatment at a higher temperature than 135° C. but below 170° C. for more than 5 hours and less than 15 hours. The invention concerns a weldable plate product of such high strength Al—Zn—Cu—Mg having a thickness of more than 50 mm and an aircraft structural member produced from such alloy.

The present invention relates to a method for producing a high strengthAl—Zn—Cu—Mg alloy with an improved corrosion resistance while at thesame time maintaining a high damage tolerance, a plate product of a highstrength Al—Zn—Cu—Mg alloy produced in accordance with the inventivemethod having a thickness of more than 50 mm and an aircraft structuralmember produced from such alloy. More specifically, the presentinvention relates to a high strength Al—Zn—Cu—Mg alloy designated by the7000-series of the international nomenclature of the AluminumAssociation for structural aeronautical applications. Even morespecifically, the present invention relates to a thick aluminum alloyproduct having improved combinations of strength, toughness andcorrosion resistance, particularly a good strength-corrosion balance.

It is known in the art to use heat treatable aluminum alloys in a numberof applications involving relatively high strength, high toughness andcorrosion resistance such as aircraft fuselages, vehicular members andother applications. Aluminum alloys AA7050 and AA7150 exhibit highstrength in T6-type tempers, see for example U.S. Pat. No. 6,315,842.Also precipitation-hardened AA7×75 alloy products exhibit high strengthvalues in the T6 temper. The T6 temper is known to enhance the strengthof the alloy, wherein the aforementioned AA7050, AA7×50 and AA7×75 alloyproducts which contain high amounts of zinc, copper and magnesium areknown for their high strength-to-weight ratios and, therefore, findapplication in particular in the aircraft industry. However, theseapplications result in exposure to a wide variety of climatic conditionsnecessitating careful control of working and ageing conditions toprovide adequate strength and resistance to corrosion, including bothstress corrosion and exfoliation.

In order to enhance resistance against stress corrosion and exfoliationas well as fracture toughness it is known to artificially over-age these7000-series alloys. When artificially aged to a T79, T76, T74 orT73-type temper their resistance to stress corrosion, exfoliationcorrosion and fracture toughness improve in the order stated (T73 beingbest and T79 being close to T6) but at some cost to strength compared tothe T6 temper condition. An acceptable temper condition is the T74-typetemper, which is a limited over-aged condition, between T73 and T76, inorder to obtain an acceptable level of tensile strength, stresscorrosion resistance, exfoliation corrosion resistance and fracturetoughness. Such a T74 temper is performed by over-ageing the aluminumalloy product at temperatures of 121° C. for 6 to 24 hours and 171° C.for about 14 hours.

Depending on the design criteria for a particular airplane componenteven small improvements in strength, toughness or corrosion resistanceresult in weight savings, which translate to fuel economy over the lifetime of the aircraft. To meet these demands several other AA7000-seriesalloys have been developed.

U.S. Pat. No. 4,954,188 discloses a method for providing a high strengthaluminum alloy characterized by improved resistance to exfoliation usingan alloy consisting of the following alloying elements, in wt. %: Zn:5.9-8.2 Cu: 1.5-3.0 Mg: 1.5-4.0 Cr: <0.04,other elements such as zirconium, manganese, iron, silicon and titaniumin total less than 0.5, the balance aluminum, working the alloy into aproduct of a pre-determined shape, solution heat treating reshapedproduct, quenching, and ageing the heat treated and quenched product toa temperature of from 132° C. to 140° C. for a period of from 6 to 30hours. The desired properties of having high strength, high toughnessand high corrosion resistance were achieved in this alloy by loweringthe ageing temperature rather than raising the temperature as taughtpreviously from, e.g., U.S. Pat. No. 3,881,966 or U.S. Pat. No.3,794,531.

It has been reported that the known precipitation-hardened aluminumalloys AA7075 and other AA7000-series alloys, in the T6 tempercondition, have not given sufficient resistance to corrosion undercertain conditions. The T7-type tempers which improve the resistance ofthe alloys to stress-corrosion cracking, however, decrease strengthsignificantly vis-à-vis the T6 condition.

U.S. Pat. No. 4,863,528 therefore discloses a method for producing animproved aluminum alloy product, the method including providing an alloyconsisting essentially of, in wt. %: Zn:  6-16 Cu: 1-3 Mg:  1.5-4.5,one or more elements selected from Zr, Cr, Mn, Ti, V, or Hf, the totalof the elements not exceeding 1.0 wt. %, the balance aluminum andincidental impurities. The aluminum alloy is solution heat-treated aftercasting, precipitation-hardened to increase its strength to a levelexceeding the as-solution heat treated strength level by about 30% ofthe difference between as-solution heat-treated strength andpeak-strength and thereafter subjected to a treatment at a sufficienttemperature or temperatures for improving its corrosion resistanceproperties. Thereafter, the alloy is again precipitation-hardened toraise its yield strength and produce a corrosion resistant alloyproduct. The ageing temperatures disclosed therein are between 170° C.and 260° C. in a range of 0.2 min. to 3 hours. The artificial ageingstep is thereby preceded and succeeded by a precipitation-hardeningstep, also known as T77 ageing. Tensile strength values were obtained ofbetween 460 MPa and 486 MPa and yield strength of 400 MPa to 434 MPa.

U.S. Pat. No. 5,035,754 discloses a heat-treating method for a highstrength aluminum alloy comprising the steps of solution heat-treatingan aluminum alloy consisting essentially of, in wt. %: Zn: 3-9 Cu: 1-3Mg:  1-6,

at least one element selected from the group consisting of Cr: 0.1-0.5Zr: 0.1-0.5 Mn:  0.2-1.0,the balance being aluminum, heating of the alloy to a temperature of alower temperature zone of 100° C. to 140° C., optionally maintaining thealloy at a temperature within the lower temperature zone for a certainduration of time, re-heating the alloy to a temperature of an uppertemperature zone of from 160° C. to 200° C., optionally maintaining thealloy at a temperature within the upper temperature zone for a secondduration of time, cooling of the alloy to a temperature of a lowertemperature zone and repeating the above mentioned steps at least twice.Such alloy improves the properties of AA7075 and AA7050 aluminum alloysby obtaining a good corrosion resistance and a high strengthcharacteristic. Some samples show a tensile strength of 57 to 62 kgf/mm²and values of the exfoliation rating of P or EA. The threshold stressvalue of the SCC-test was more than 50 kgf/mm².

EP-0377779 discloses a process for producing an alloy for sheet or thinplate applications in the field of aerospace such as upper-wing memberswith high toughness and good corrosion properties which comprises thesteps of working a body having a composition consisting of, in wt. %:Zn: 7.6-8.4 Cu: 2.2-2.6 Mg:  1.8-2.1,

and one or more elements selected from Zr:  0.5-0.2 Mn: 0.05-0.4 V:0.03-0.2 Hf:  0.03-0.5,the total of the elements not exceeding 0.6 wt. %, the balance aluminumplus incidental impurities, solution heat treating and quenching theproduct and artificially ageing the product by either heating theproduct three times in a row to one or more temperatures from 79° C. to163° C. or heating such product first to one or more temperatures from79° C. to 141° C. for two hours or more or heating the product to one ormore temperatures from 148° C. to 174° C. These products show animproved exfoliation corrosion resistance of “EB” or better with about15% greater yield strength than similar sized AA7×50 counter parts inthe T76-temper condition. They still have at least about 5% greaterstrength than their similarly sized AA7×50-T77 counter-part.

U.S. Pat. No. 5,312,498 discloses another method for producing analuminum-based alloy product having improved exfoliation resistance andfracture toughness with balanced zinc, copper and magnesium levels suchthat there is no excess of copper and magnesium. The method of producingthe aluminum-based alloy product utilizes either a one-step or two-stepageing process in conjunction with the stoichiometric balancing ofcopper, magnesium and zinc. A two-step ageing sequence is disclosedwherein the alloy is first aged at about 121° C. for about 9 hoursfollowed by a second ageing step at about 157° C. for about 10 to 16hours followed by air cooling. Such ageing method is directed to thinplate or sheet products that are used for lower-wing skin applicationsor fuselage skin.

There is, however, a demand in the fields of aeronautics to provide highstrength AA7000-series alloys with a cross-sectional thickness of morethan 50 mm for e.g. spars or bars of wings and upper-wing skinapplications with the above mentioned specific mechanical propertiessuch as high strength, high toughness and good corrosion properties suchas resistance to stress corrosion or resistance to exfoliationcorrosion. These parts such as spars of wings for aircraft are typicallymanufactured from a plate product via machining operations wherein thematerial property is a compression yield strength in the L-direction atS/4 of at least 475 MPa, an ultimate tensile strength of at least 510MPa and an ST (short transverse) elongation at S/2 of at least 3.0%.

EP-1158068A1 discloses a heat-treatable aluminum alloy for producingthick products having a thickness of more than 12 mm, the alloy is anAl—Zn—Cu—Mg alloy with the following composition, in wt. %: Zn:  4-10Cu:   1-3.5 Mg: 1-4 Cr: <0.3 Zr: <0.3 Si: <0.5 Fe: <0.5other elements <0.05 each and <0.15 in total, balance aluminum. It isdisclosed that it was found that for thick products with an onlyslightly recrystallized microstructure, a high as-cast grain size couldlead to a specific microstructure of the transformed and heat-treatedproduct which has a beneficial effect on the toughness with no reductionin strength or other properties. It is therefore described to cast thealloy in the form of a rolling, forging or extrusion ingot such that theas-cast grain size is kept between 300 and 800 μm.

It is therefore an object of the present invention to provide animproved method for producing a high strength Al—Zn—Cu—Mg alloy forthick plate products with an improved fatigue crack growth resistanceand a high damage tolerance which has the aforementioned properties of acompression yield strength (in L-direction at S/4) of at least 475 MPa,an ultimate tensile strength of at least 510 MPa and an ST elongation atS/2 of at least 3.0%.

It is another object of the invention to obtain an AA7000-seriesaluminum alloy, which exhibits strength in the range of T6-type tempersand toughness and corrosion resistance properties in the range ofT73-type tempers.

It is furthermore an object of the present invention to obtain a thickplate alloy, which can be used to produce structural parts of aircraftsuch as spars of wings with high strength levels and good corrosionresistance properties.

The present invention meets these objects by the characterizing featuresof claim 1. Further preferred embodiments are described and specifiedwithin the sub-claims.

According to the invention there is disclosed a method for producing ahigh strength Al—Zn—Cu—Mg alloy with an improved fatigue crack growthresistance and a high damage tolerance, comprising the steps of:

a) casting an ingot with the following composition (in weight percent):Zn: 5.5-9.5 Cu: 1.5-3.5 Mg: 1.5-3.5 Mn: <0.25 Zr: <0.25, preferably0.06-0.16 Cr: <0.10 Fe: <0.25, preferably <0.15 Si: <0.25, preferably<0.10 Ti: <0.10

-   -    Hf and/or V <0.25, and    -    other elements each less than 0.05 and less than 0.15 in total,        balance aluminum,    -   b) homogenizing and/or pre-heating the ingot after casting,    -   c) hot working the ingot, preferably by means of rolling, and        optionally cold working, preferably by means of rolling, into a        worked product of more than 50 mm thickness,    -   d) solution heat-treating,    -   e) quenching the solution heat treated product, and artificially        ageing the worked and heat-treated product, wherein the ageing        step comprises a first heat treatment at a temperature in a        range of 105° C. to 135° C. for more than 2 hours and less than        8 hours and a second heat treatment at a higher temperature than        135° C. but below 170° C. for more than 5 hours and less than 15        hours to achieve a product with a compression yield strength in        L-direction at S/4 of at least 475 MPa, an ultimate tensile        strength of at least 510 MPa and an ST elongation at S/2 of at        least 3.0%.

The above mentioned combination of chemistry and ageing practiceexhibits very high strength levels, very good exfoliation resistance andhigh stress corrosion resistance for thick plate products with thicknessof more than 50 mm. Specifically, the two-step ageing practice of thepresent invention utilizes a first heat treatment for 2 to 5 hours, attemperatures in the range of 115° C. to 125° C., preferably about 4hours at 120° C. and a second heat treatment for 5 to 15 hours, attemperatures in the range of 155° C. to 169° C., preferably for about 13hours at temperatures between 161° C. to 167° C.

It will be immediately apparent to the skilled person that in the methodaccording to this invention, that after quenching of the solution heattreated product and before the artificial ageing practice, the productmay optionally be stretched or compressed or otherwise cold worked torelieve stresses as known in the art.

Preferred amounts (in wt. %) of magnesium are in a range of 1.5 to 2.5,preferably in a range of 1.6 to 2.3, and more preferably in the range of1.90 to 2.10. Preferred amounts (in wt. %) of copper are in a range of1.5 to 2.5, preferably in a range of 1.6 to 2.3, and more preferably inthe range of 1.85 to 2.10. Preferred amounts (in wt. %) of zinc are in arange of 5.9 to 6.2 or in a range of 6.8 to 7.1 or in a range of 7.8 to8.1.

Copper and magnesium are important elements for adding amongst othersstrength to the alloy. The preferred range of copper and magnesium isabove 1.6 wt. % and lower than 2.3 wt. % since too low amounts ofmagnesium and copper result in a decrease of strength while too highamounts of magnesium and copper result in a lower corrosion performanceand problems with the weldability of the alloy product. In order toachieve a compromise in strength, toughness and corrosion performanceeach of copper and magnesium amounts (in weight %) of between 1.6 and2.3, with preferred narrower ranges set out above and in the claims,have been found to give a good balance for thick alloy products. If theamounts of copper and magnesium are chosen too high the propertiesrelating to toughness, stress corrosion and elongation will drop,especially for thicker products.

Furthermore, it has been found that the balance of copper and magnesiumto zinc, especially the balance of magnesium to zinc is of importance.Depending on the amount of zinc the amount (in wt. %) of magnesium ispreferably in between 2.4-0.1[Zn] and 1.5+0.1[Zn]. That means that theamount of magnesium depends on the chosen amount of zinc. With an amountof approximately 6 wt. % Zn the amount (in wt. %) of magnesium isbetween 1.8 and 2.1, when Zn is approximately 7% the amount of magnesiumis between 1.7 and 2.2 and if Zn is approximately 8% the amount ofmagnesium is between 1.6 and 2.3.

With the method according to the present invention and the chosenbalance of copper, magnesium and zinc it is possible to obtain ahomogenized and/or pre-heated ingot after casting which is hot-workedand optionally cold-worked into a worked product of preferably more than60 mm thickness, more preferably in a range of 110 mm to 160 mm and evenup to 220 mm thickness with an improved corrosion performance which isat least as good as achievable with the T77 ageing method but lesscomplicated than the so-called three-step-ageing temper T77.

The alloy of the present invention is preferably selected from the groupconsisting of AA7010, AA7×50, AA7040, AA7020, AA7×75, AA7349 or AA7×55or AA7×85, preferably AA7055 or AA7085.

According to the invention there is disclosed a plate product of highstrength alloy-zinc-copper-magnesium-alloy produced in accordance with amethod as defined above having a thickness of more than 50 mm,preferably 100 mm to 220 mm. Such plate product is preferably a part ofan aircraft such as a bar or a spar of a wing. Most preferably, theplate product according to the present invention is an upper-wing memberof an aircraft.

EXAMPLES

The foregoing and other features and advantages of the alloys accordingto the invention will become readily apparent from the followingdetailed description of preferred embodiments.

On an industrial scale 7 different aluminum alloys have been cast intoingots having the following chemical composition as set out in Table 1.TABLE 1 Chemical composition of thick plate alloys, in wt. %, balancealuminum and inevitable impurities, Fe = 0.08 and Si = 0.04, and Zr =0.10, Alloys 1 to 5 with Mn = 0.02 and Alloys 6 and 7 with Mn = 0.08.Alloying Element Alloy Cu Mg Zn Zr 1 2.16 2.04 6.18 0.11 2 2.10 2.006.10 0.10 3 2.14 2.04 6.12 0.10 4 1.91 2.13 6.86 0.11 5 2.20 2.30 6.900.10 6 2.23 2.50 7.80 0.10 7 1.82 2.18 8.04 0.10

Full scale ingots have been sawn from the ingot slices, homogenized for12 hours 470° C. and for 24 hours at 475° C., pre-heated for 5 hours at410° C. and hot-rolled to a thickness of various gauges as identified inTable 2. Thereafter, the plates were solution heat treated for 4 hoursat 475° C. with subsequent quenching and a two-step ageing process,first for 4 hours at 120° C. and second for 13 hours at 165° C.

The alloys mentioned in Table 1 were examined with regard to variousplate thicknesses as identified in Table 2. TABLE 2 Overview ofstrength, elongation and exfoliation properties of different thicknessof the alloys of Table 1 (S/2 = mid-thickness; S/4 = quarter-thickness);EXCO testing at S/10 according to ASTM G34, samples shown for EA-EDclassification. Rm-L Plate thickness Rp-L (MPa) (MPa) A-(ST) (%) (mm)Alloy S/4 S/4 S/2 EXCO 63.5 1 553 590 6 EC 110 2 503 553 4 EA 152 3 495537 5 EA 152  3* 480 528 5 EA 63.5 4 570 604 3 EC 110 5 515 550 2 EA 1106 510 565 2 EA 152 7 476 529 3 EA*aged at 120° for 5 hours and subsequently at 165° C. at 15 hours.

As shown in Table 2 the alloys of Table 1 show good compression yieldstrength (“Rp”) in the L-direction of more than 476 MPa, most of themmore than 500 MPa while the ultimate tensile strength (“Rm”) in theL-direction is above 529 MPa for alloys and thickness, one example evenabove 600 MPa for 63.5 mm. The ST-elongation at position S/2 of all buttwo alloys is 3% or above, even up to 6%.

The exfoliation properties are EA or EC. The exfoliation testing wasdone in accordance to ASTM G34 at S/10 position. The exfoliationproperties are similar for similar ageing steps as shown in Table 3 butsurprisingly deteriorate if the first heat treatment is longer and thesecond heat treatment is shorter. TABLE 3 Exfoliation properties(“EXCO”) of selected alloys of Table 1 according to ASTM G34 (“—” meansnot measured). 6 h/120° C. + 5 h/120° C. + 4 h/120° C. + Alloy Thickness6 h/155° C. 12 h/155° C. 13 h/165° C. 1 63.5 EC — EC 3 110 — EA EA 563.5 EC — — 5 110 EC EA EA 6 110 ED EA EA 7 63.5 EC — EA

Alloy 4 has been tested with a plate thickness of 110 mm. The toughnessand elongation are shown in Table 4. TABLE 4 Toughness and elongationproperties of selected alloys of Table 1, all plates of 110 mmthickness, ageing according to a two-step method, first heat treatmentat 120° C. for 4 hours, second heat treatment at 165° C. for 13 hours,alloy 5 with a copper content of 2.25; K_(IC) measured according to normASTM E399-90 C(T) specimens, thickness of 38.1 mm (1.5″) for SL, SLsamples taken from mid-thickness (S/2). Rp A K_(IC) Alloy (S/2, ST)(S/2, ST) (S/2, SL) 1 465 5 26.9 3 461 5 26.8 4 465 5 27.1 5 453 2 24.16 472 1 19.5 7 482 3 26.4

All above mentioned alloys showed an exfoliation rating of EA for theselected plate thickness of 110 mm.

Finally, the stress corrosion properties (“SCC”) were examined. First,alloys 1 and 4 were tested with thickness of 152 mm. Two differentageing procedures were selected in accordance with Table 5. The loadlevel was 172 MPa. The test direction is S-L. Samples were taken fromthe S/2 position. Table 5 shows the number of days till failure wasgiven. After 30 days the test was terminated. “NF” means no failureafter 30 days, “30” means failure after 30 days. In total at least threespecimens are tested per variant. The test was done in accordance withASTM G47. TABLE 5 SCC-properties for thickness of 152 mm for two alloys.Alloy 5 h/120° C. + 12 h/165° C. 4 h/120° C. + 15 h/165° C. 1 NF, NF, NFNF, NF, NF 4 30, NF, NF NF, NF, NF

Finally, 5 other alloys were tested with regard to the stress corrosionproperties by using plates of a thickness of 125 mm. Samples were takenfrom the S-L direction at a load level of 180 MPa. Table 6 shows thechemistry and the results of those alloys with regard to the stresscorrosion properties. TABLE 6 SCC-properties of S-L specimens having athickness of 125 mm, Fe = 0.08, Si = 0.04, and Zr = 0.10. Alloy Cu Mg Zn4 h/120° C. + 13 h/165° C. A 1.7 1.8 7.4 NF, NF, NF B 2.3 1.8 7.5 NF,NF, NF C 2.25 2.5 7.65 15, NF, NF D 1.8 2.45 8.0 15, 20, NF E 2.3 2.48.1 20, 25, NF

As can be seen from Table 6 the toughness of the inventive alloy iscontrolled by the copper and magnesium levels while zinc has aninfluence in particular on the tensile properties. The preferred balanceof each of copper and magnesium is in between 1.6 and 2.0 wt. %.

Having now fully described the invention, it will be apparent to one ofordinary skill in the art that many changes and modifications can bemade without departing from the scope or spirit of the invention asherein described.

1. Method for producing a high strength Al—Zn—Cu—Mg alloy with a highdamage tolerance and an improved corrosion resistance, comprising thesteps of: a) casting an ingot with the following composition (in weightpercent): Zn 5.5 to 9.5 Cu 1.5 to 3.5 Mg 1.5 to 3.5 Mn <0.25 Zr <0.25 Cr<0.10 Fe <0.25 Si <0.25 Ti <0.10

 Hf and/or V<0.25  other elements each less than 0.05 and less than 0.15in total, balance aluminum, b) homogenizing and/or pre-heating the ingotafter casting, c) hot-working the ingot and optionally cold working intoa worked product of more than 50 mm thickness, d) solutionheat-treating, e) quenching the solution heat treated product, and f)artificially ageing the worked and heat-treated product, wherein theageing step comprises a first heat treatment at a temperature in a rangeof 105° C. to 135° C. for more than 2 hours and less than 8 hours and asecond heat treatment at a higher temperature than 135° C. but below170° C. for more than 5 hours and less than 15 hours to achieve aproduct with a compression yield strength in L-direction at S/4 of atleast 475 MPa, an ultimate tensile strength of at least 510 MPa and anST elongation at S/2 of at least 3.0%.
 2. Method according to claim 1,wherein the ageing step consists of two heat treatments, the first heattreatment is performed for 2 to 5 hours at temperatures in the range of105° C. to 135° C., and the second heat treatment is performed for 5 to15 hours at temperatures in the range of 155° C. to 169° C.
 3. Methodaccording to claim 1, wherein the first heat treatment is performed attemperatures in the range 115° C. to 125° C.
 4. Method according toclaim 1, wherein the first heat treatment is performed for 2 to 5 hoursat about 120° C.
 5. Method according to claim 1, wherein the second heattreatment is performed at temperatures in the range 161° C. to 167° C.6. Method according to claim 1, wherein the second heat treatment isperformed for about 13 hours.
 7. Method according to claim 1, whereinthe improved corrosion resistance has exfoliation properties (“EXCO”) ofEB or better according to ASTM G34.
 8. Method according to claim 1,wherein the amount of Mg is in a range of 1.5 to 2.5.
 9. Methodaccording to claim 8, wherein the amount of Mg is in a range of 1.6 to2.3.
 10. Method according to claim 8, wherein the amount of Mg is in arange of 1.90 to 2.10.
 11. Method according to claim 1, wherein theamount of Cu is in a range of 1.5 to 2.5.
 12. Method according to claim11, wherein the amount of Cu is in a range of 1.6 to 2.3.
 13. Methodaccording to claim 11, wherein the amount of Cu is in a range of 1.85 to2.1
 0. 14. Method according to claim 1, wherein the amount of Mg dependson the amount of Zn as follows: [Mg] is in between 2.4-0.1[Zn] and1.5+0.1[Zn].
 15. Method according to claim 1, wherein the amount of Znis in a range of 5.9 to 6.2.
 16. Method according to claim 1, whereinthe amount of Zn is in a range of 6.8 to 7.1.
 17. Method according toclaim 1, wherein the amount of Zn is in a range of 7.8 to 8.1. 18.Method according to claim 1, wherein the amount of Fe is less than 0.15.19. Method according to claim 1, wherein the amount of Fe is 0.08 orless.
 20. Method according to claim 1, wherein the amount of Si is lessthan 0.10.
 21. Method according to claim 1, wherein the amount of Si is0.04 or less.
 22. Method according to claim 1, wherein the amount of Zris in a range of 0.06 to 0.16.
 23. Method according to claim 1, whereinthe amount of Mn is 0.08 or less.
 24. Method according to claim 1,wherein the amount of Mn is 0.02 or less.
 25. Method according to claim1, wherein after the step e) the quenched solution heat-treated productis stretched or compressed or otherwise cold worked to relieve stressesprior to the ageing practice of the step f).
 26. Method according toclaim 1, wherein the product is hot-worked by means of rolling. 27.Method according to claim 1, wherein the product is cold-worked by meansof rolling.
 28. Method according to claim 1, wherein after homogenizingand/or pre-heating the ingot after casting, hot working the ingot andoptionally cold working the ingot into a worked product of more than 60mm.
 29. Method according to claim 1, wherein after homogenizing and/orpre-heating the ingot after casting, hot working the ingot andoptionally cold working the ingot into a worked product of more than 110mm.
 30. Method according to claim 1, wherein after homogenizing and/orpre-heating the ingot after casting, hot working the ingot andoptionally cold working the ingot into a worked product of not more than220 mm.
 31. Method according to claim 1, wherein after homogenizingand/or pre-heating the ingot after casting, hot working the ingot andoptionally cold working the ingot into a worked product of not more than160 mm.
 32. Method according to claim 1, wherein said high strengthAl—Zn—Cu—Mg alloy is selected from the group consisting of AA7010,AA7×50, AA7040, AA7020, AA7×75, AA7349, AA7×55, and AA7×85.
 33. A plateproduct of high strength Al—Zn—Cu—Mg alloy produced in accordance with amethod as defined in claim 1 and having a thickness of more than 50 mm.34. A plate product of high strength Al—Zn—Cu—Mg alloy produced inaccordance with a method as defined in claim 1 and having a thickness ina range of 110 to 220
 35. A plate product of high strength Al—Zn—Cu—Mgalloy produced in accordance with a method as defined in claim 1 andhaving a thickness of more than 60 mm.
 36. A plate product according toclaim 33, wherein said plate product is a structural member of anaircraft.
 37. A plate product according to claim 34, wherein said plateproduct is a structural member of an aircraft.
 38. A plate productaccording to claim 33, wherein said plate product is a bar or a spar ofa wing of an aircraft.
 39. A plate product according to claim 34,wherein said plate product is an upper-wing member of an aircraft. 40.An aircraft structural member produced from a high strength Al—Zn—Cu—Mgalloy produced in accordance with a method as defined in claim
 1. 41. Anaircraft structural member having a thickness of at least 50 mm andmanufactured from a rolled product made of an alloy with a composition,consisting of, in % by weight: Zn 5.5to9.5 Cu 1.5 to 3.5 Mg 1.5 to 3.5Mn<0.25 Zr<0.25 Cr<0.10 Fe<0.25 Si<0.25 Ti<0.10 Hf and/or V<0.25 otherelements each less than 0.05 and less than 0.15 in total, balancealuminum, and treated by solution heat treating, quenching, and ageingpractice consisting essentially of a first heat treatment at atemperature in a range of 105° C. to 135° C. for more than 2 hours andless than 8 hours and a second heat treatment at a higher temperaturethan 135° C. but below 170° C. for more than 5 hours and less than 15hours, the product having a compression yield strength in L-direction atS/4 of at least 475 MPa, an ultimate tensile strength of at least 510MPa and an ST elongation at S/2 of at least 3.0%.
 42. An aircraftstructural member according to claim 41, wherein the aircraft structuralmember has a thickness in a range of 50 to 220 mm.
 43. An aircraftstructural member according to claim 41, wherein the aircraft structuralmember has a thickness in a range of 60 to 160 mm.
 44. An aircraftstructural member according to claim 41, wherein the aircraft structuralmember has a thickness in a range of 110 to 160 mm.
 45. An aircraftstructural member according to claim 41, forming a part of an aircraftupper wing.
 46. An aircraft structural member according to claim 41,forming a spar of an aircraft wing.
 47. An aircraft structural memberaccording to claim 41, forming a bar of an aircraft wing.
 48. Anaircraft structural member according to claim 41, obtained by machining.49. An aircraft structural member according to claim 41, wherein theageing practice consists of two heat treatments, the first heattreatment is performed for 2 to 5 hours at temperatures in the range of105° C. to 135° C., and the second heat treatment is performed for 5 to15 hours at temperatures in the range of 155° C. to 169° C.
 50. Anaircraft structural member according to claim 41, wherein the first heattreatment of the ageing practice is performed at temperatures in therange 115° C. to 125° C.
 51. An aircraft structural member according toclaim 41, wherein the first heat treatment of the ageing practice isperformed for 2 to 5 hours at about 120° C.
 52. An aircraft structuralmember according to claim 41, wherein the second heat treatment of theageing practice is performed at temperatures in the range 161° C. to167° C.
 53. An aircraft structural member according to claim 41, whereinthe second heat treatment of the ageing practice is performed for about13 hours.
 54. An aircraft structural member according to claim 41,wherein the improved corrosion resistance has exfoliation properties(“EXCO”) of EB or better according to ASTM G34.
 55. An aircraftstructural member according to claim 40, wherein in the amount of Mg isin a range of 1.5 to 2.5.
 56. An aircraft structural member according toclaim 55, wherein in the amount of Mg is in a range of 1.6 to 2.3. 57.An aircraft structural member according to claim 55, wherein in theamount of Mg is in a range of 1.90 to 2.10.
 58. An aircraft structuralmember according to claim 41, wherein the amount of Cu is in a range of1.5 to 2.5.
 59. An aircraft structural member according to claim 58,wherein the amount of Cu is in a range of 1.6 to 2.3.
 60. An aircraftstructural member according to claim 58, wherein the amount of Cu is ina range of 1.85 to 2.10.
 61. An aircraft structural member according toclaim 41, wherein the amount of Mg depends on the amount of Zn asfollows: [Mg] is in between 2.4-0.1[Zn] and 1.5+0.1[Zn].
 62. An aircraftstructural member according to claim 41, wherein the amount of Zn is ina range of 5.9 to 6.2.
 63. An aircraft structural member according toclaim 41, wherein the amount of Zn is in a range of 6.8 to 7.1.
 64. Anaircraft structural member according to claim 41, wherein the amount ofZn is in a range of 7.8 to 8.1.
 65. An aircraft structural memberaccording to claim 41, wherein the amount of Fe is less than 0.15. 66.An aircraft structural member according to claim 41, wherein the amountof Si is less than 0.10.
 67. An aircraft structural member according toclaim 41, wherein the amount of Zr is in a range of 0.06 to 0.16.
 68. Anaircraft structural member according to claim 41, wherein the amount ofMn is in a range of 0.08 or less.
 69. An aircraft structural memberaccording to claim 41, wherein after the quenching following thesolution heat-treating the product is stretched or compressed orotherwise cold worked to relieve stresses prior to the ageing practice.